This invention relates generally to gas turbine engines, and more specifically to rotor blades used with axial gas turbine engines.
Gas turbine engines include a rotor assembly including a row of rotor blades. The blades extend radially outward from a platform that extends between an airfoil portion of the blade and a dovetail portion of the blade. The platform defines a portion of the gas flow path through the engine, and the dovetail couples each rotor blade to the rotor disk. More specifically, each rotor blade extends radially outward from the platform to a tip. A plurality of static shrouds abut together to form flowpath casing that extends circumferentially around the rotor blade assembly, such that a tip clearance is defined between each respective rotor blade tip and the casing or shroud. The tip clearance is tailored to be a minimum, yet is sized large enough to facilitate rub-free engine operation through the range of available engine operating conditions.
During operation, tip leakage across the rotor blade tips may limit the performance and stability of the rotor assembly. To facilitate increasing an efficiency and a stable flow range (a stall margin) at a given clearance for the rotor assembly at least some known rotor assemblies, grooves are machined into the flowpath casing above the rotor tips to facilitate increasing pressure rise and stability of the airflow. Such grooves, known as casing treatments, may have an efficiency penalty that increases with their effectiveness in delaying stall. Additionally, such casing treatments may not reduce the sensitivity of performance and stall margin that may be caused with increased tip clearance levels. To prevent inducing fatigue stresses into the shroud, the shrouds are often fabricated from stronger and thicker materials, and as such, the casing treatments may also increase an overall weight of the rotor assembly.
In one aspect of the invention, an airfoil for a gas turbine engine is provided. The airfoil includes a leading edge, a trailing edge, a tip plate, a first sidewall, and a second sidewall. The first sidewall extends in radial span between an airfoil root and the tip plate, and defines a pressure side of the airfoil. The second sidewall is connected to the first sidewall at the leading and trailing edges, and extends in radial span between the airfoil root and the tip plate to define a suction side of the airfoil. The tip plate includes at least one groove that extends substantially between the first and second sidewalls.
In another aspect, a method for fabricating a rotor blade for a gas turbine engine is provided. More specifically, the method facilitates improving an efficiency of the rotor blade. The method includes casting a rotor blade to include a leading edge, a trailing edge, a first sidewall, and a second sidewall, wherein the first and second sidewalls are connected chordwise at the leading and trailing edges, and extend radially between a blade root and a blade tip plate, and forming at least one groove in the tip plate that extends substantially between the first and second sidewalls.
In a further aspect, a gas turbine engine including a plurality of rotor blades is provided. Each of the rotor blades includes an airfoil including a leading edge, a trailing edge, a first sidewall, a second sidewall, and a tip plate. The airfoil first and second sidewalls are connected chordwise at the leading and trailing edges. The first and second sidewalls extend radially from a blade root to the tip plate, and the tip plate includes a groove that extends substantially between the airfoil first and second sidewalls. The groove is for transferring fluid from the first sidewall to the second sidewall.